Low noise compressor rotor for geared turbofan engine

ABSTRACT

A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor, and having a gear reduction ratio of greater than 2.5:1. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor: (number of blades×rotational speed)/60 sec≧5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.14/591,975, filed Jan. 8, 2015, which is a continuation-in-part of U.S.patent application Ser. No. 14/144,710, filed Dec. 31, 2013, which is acontinuation of U.S. patent application Ser. No. 14/016,436, filed Sep.3, 2013, now U.S. Pat. No. 8,714,913, issued May 6, 2014, which is acontinuation of U.S. patent application Ser. No. 13/630,276, filed Sep.28, 2012, now U.S. Pat. No. 8,632,301, issued Jan. 21, 2014.

BACKGROUND

This application relates to the design of a gas turbine engine rotorwhich can be operated to produce noise that is less sensitive to humanhearing.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor. The air is compressed in the compressor anddelivered downstream into a combustor section where it was mixed withfuel and ignited. Products of this combustion pass downstream overturbine rotors, driving the turbine rotors to rotate.

Typically, there is a high pressure turbine rotor, and a low pressureturbine rotor. Each of the turbine rotors include a number of rows ofturbine blades which rotate with the rotor. Interspersed between therows of turbine blades are vanes.

The high pressure turbine rotor has typically driven a high pressurecompressor rotor, and the low pressure turbine rotor has typicallydriven a low pressure compressor rotor. Each of the compressor rotorsalso include a number of compressor blades which rotate with the rotors.There are also vanes interspersed between the rows of compressor blades.

The low pressure turbine or compressor can be a significant noisesource, as noise is produced by fluid dynamic interaction between theblade rows and the vane rows. These interactions produce tones at ablade passage frequency of each of the low pressure turbine rotors, thelow pressure compressor rotors, and their harmonics.

The noise can often be in a frequency range that is very sensitive tohumans. To mitigate this problem, in the past, a vane-to-blade ratio hasbeen controlled to be above a certain number. As an example, avane-to-blade ratio may be selected to be 1.5 or greater, to prevent afundamental blade passage tone from propagating to the far field. Thisis known as “cut-off.”

However, acoustically cut-off designs may come at the expense ofincreased weight and reduced aerodynamic efficiency. Stated another way,by limiting the designer to a particular vane to blade ratio, thedesigner may be restricted from selecting such a ratio based upon othercharacteristics of the intended engine.

Historically, the low pressure turbine has driven both a low pressurecompressor section and a fan section. More recently, a gear reductionhas been provided such that the fan and low pressure compressor can bedriven at distinct speeds.

SUMMARY OF THE INVENTION

A gas turbine engine according to an example of the present disclosureincludes a fan, a turbine section that has a fan drive turbine, and acompressor section that a first compressor. A gear reduction ispositioned between the fan on one side and the fan drive turbine onanother side. The first compressor has a number of compressor blades inat least one of a plurality of rows of the first compressor, and theblades are rotatable at least some of the time at a rotational speed inoperation. The number of compressor blades in the at least one row andthe rotational speed are such that the following formula holds true forthe at least one row of the first compressor: (the number of blades×therotational speed)/60 sec≧5500 Hz. The rotational speed is an approachspeed in revolutions per minute, taken at an approach certificationpoint as defined in Part 36 of the Federal Airworthiness Regulations,and a pressure ratio across the fan drive turbine being greater than 5.

In a further embodiment of any of the foregoing embodiments, the gasturbine engine is rated to produce 15,000 pounds of thrust or more.

In a further embodiment of any of the foregoing embodiments, the gearreduction has a gear ratio of greater than 2.3.

In a further embodiment of any of the foregoing embodiments, the fandelivers air into a bypass duct, and a portion of air into thecompressor section, with a bypass ratio defined as the volume of airdelivered into the bypass duct compared to the volume of air deliveredinto the compressor section. The bypass ratio is greater than 6.

In a further embodiment of any of the foregoing embodiments, the gearreduction has a gear ratio of greater than about 2.5.

In a further embodiment of any of the foregoing embodiments, the bypassratio is greater than 10.

In a further embodiment of any of the foregoing embodiments, the fancomprises at least one fan blade, with a low fan pressure ratio of lessthan 1.45. The low fan pressure ratio is measured across the fan bladealone.

In a further embodiment of any of the foregoing embodiments, the turbinesection includes a higher pressure turbine and a lower pressure turbine,and the fan drive turbine is the lower pressure turbine.

In a further embodiment of any of the foregoing embodiments, the firstcompressor is a lower pressure compressor, and the higher pressureturbine drives a higher pressure compressor.

In a further embodiment of any of the foregoing embodiments, the gearreduction is positioned intermediate the fan drive turbine and the firstcompressor.

In a further embodiment of any of the foregoing embodiments, the formuladoes not hold true for each of the rows of the first compressor.

In a further embodiment of any of the foregoing embodiments, the formulaholds true for a majority of the rows of the first compressor.

In a further embodiment of any of the foregoing embodiments, the gearreduction is positioned intermediate the fan and a compressor driven bythe fan drive turbine.

In a further embodiment of any of the foregoing embodiments, the fan hasa low corrected fan tip speed of less than 1150 ft/second.

In a further embodiment of any of the foregoing embodiments, the formularesults in a number greater than or equal to 6000 Hz.

A method of designing a gas turbine engine according to an example ofthe present disclosure includes the steps of including a turbine sectionthat has a first turbine to drive a first compressor and a fan turbinefor driving a fan through a gear reduction, with the gear reductionpositioned between the fan on one side and the fan turbine on anotherside, and selecting a combination of a number of blades in at least onerow of the first compressor and a rotational speed of the firstcompressor in operation, for producing noise frequencies that are ofless concern to humans. The selecting step includes determining thenumber of blades in the at least one row of the first compressor anddetermining a rotational speed of the first compressor in operation suchthat the following formula holds true for the at least one row of thefirst compressor: (the number of blades×the rotational speed)/60sec≧5500 Hz. The rotational speed is an approach speed in revolutionsper minute, taken at an approach certification point as defined in Part36 of the Federal Airworthiness Regulations, and a pressure ratio acrossthe fan turbine being greater than 5.

In a further embodiment of any of the foregoing embodiments, the gasturbine engine is rated to produce 15,000 pounds of thrust or more.

In a further embodiment of any of the foregoing embodiments, the turbinesection includes a higher pressure turbine and a lower pressure turbine.The fan turbine is the lower pressure turbine.

In a further embodiment of any of the foregoing embodiments, the firstcompressor is a lower pressure compressor, and the higher pressureturbine driving a higher pressure compressor.

In a further embodiment of any of the foregoing embodiments, the fancomprises at least one fan blade, with a low fan pressure ratio of lessthan 1.45. The low fan pressure ratio is measured across the fan bladealone. The fan delivers air into a bypass duct, and a portion of airinto the first compressor, with a bypass ratio defined as the volume ofair delivered into the bypass duct compared to the volume of airdelivered into the first compressor. The bypass ratio is greater than10. The fan has a low corrected fan tip speed of less than 1150ft/second.

In a further embodiment of any of the foregoing embodiments, the firstturbine and the fan turbine are provided by a single rotor.

In a further embodiment of any of the foregoing embodiments, the formuladoes not hold true for each of the rows of the first compressor.

In a further embodiment of any of the foregoing embodiments, the formulaholds true for a majority of the rows of the first compressor.

In a further embodiment of any of the foregoing embodiments, the formularesults in a number greater than or equal to 6000 Hz.

A gas turbine engine according to an example of the present disclosureincludes a fan, a turbine section that has a fan drive turbine, and acompressor section that has a first compressor. A gear reduction ispositioned between the fan on one side and the fan drive turbine onanother side. The first compressor has a number of compressor blades inat least one of a plurality of rows of the first compressor, and theblades are rotatable at least some of the time at a rotational speed inoperation. The number of compressor blades in the at least one row andthe rotational speed is such that the following formula holds true forthe at least one row of the first compressor: 5500 Hz≦(the number ofblades×the rotational speed)/60 sec≦6000 Hz. The rotational speed is anapproach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations. A pressure ratio across the fan drive turbine is greaterthan 5.

In a further embodiment of any of the foregoing embodiments, the turbinesection includes a higher pressure turbine and a lower pressure turbine,and the fan drive turbine being the lower pressure turbine.

In a further embodiment of any of the foregoing embodiments, the formuladoes not hold true for each of the rows of the first compressor.

In a further embodiment of any of the foregoing embodiments, the gearreduction has a gear ratio of greater than 2.5.

In a further embodiment of any of the foregoing embodiments, the fandelivers air into a bypass duct, and a portion of air into thecompressor section, with a bypass ratio defined as the volume of airdelivered into the bypass duct compared to the volume of air deliveredinto the compressor section. The bypass ratio is greater than 10. Thefan has a low corrected fan tip speed less than 1150 ft/second.

In a further embodiment of any of the foregoing embodiments, the formulaholds true for a majority of the rows of the first compressor.

In a featured embodiment, a gas turbine engine comprises a fan and aturbine section having a fan drive turbine rotor, and a compressorrotor. A gear reduction effects a reduction in the speed of the fanrelative to an input speed from the fan drive turbine rotor. Thecompressor rotor has a number of compressor blades in at least one of aplurality of rows of the compressor rotor. The blades operate at leastsome of the time at a rotational speed. The number of compressor bladesin the at least one row and the rotational speed is such that thefollowing formula holds true for the at least one row of the compressorrotor: (the number of blades×the rotational speed)/60 s≧5500 Hz. Therotational speed is in revolutions per minute.

In another embodiment according to the previous embodiment, the formularesults in a number greater than or equal to about 6000 Hz.

In another embodiment according to any of the previous embodiments, thegas turbine engine is rated to produce about 15,000 pounds of thrust ormore.

In another embodiment according to any of the previous embodiments, theformula holds true for the majority of the blade rows of the compressorrotor.

In another embodiment according to any of the previous embodiments, thegas turbine engine is rated to produce about 15,000 pounds of thrust ormore.

In another embodiment according to any of the previous embodiments, thegear reduction has a gear ratio of greater than about 2.3.

In another embodiment according to any of the previous embodiments, thegear reduction has a gear ratio of greater than about 2.5.

In another embodiment according to any of the previous embodiments, thefan delivers air into a bypass duct, and a portion of air into thecompressor rotor. A bypass ratio is defined as the volume of airdelivered into the bypass duct compared to the volume of air deliveredinto the compressor rotor, and the bypass ratio being greater than about6.

In another embodiment according to any of the previous embodiments, thebypass ratio is greater than about 10.

In another embodiment according to any of the previous embodiments, theformula results in a number greater than or equal to about 6000 Hz.

In another embodiment according to any of the previous embodiments, therotational speed being an approach speed.

In another embodiment according to any of the previous embodiments, theturbine section includes a higher pressure turbine rotor and a lowerpressure turbine rotor. The fan drive turbine rotor is the lowerpressure turbine rotor.

In another embodiment according to any of the previous embodiments, thecompressor rotor is a lower pressure compressor rotor, and the higherpressure turbine rotor drives a higher pressure compressor rotor.

In another embodiment according to any of the previous embodiments,there are three turbine rotors. The fan drive rotor turbine drives thefan. A second and third turbine rotor each drive respective compressorrotors of the compressor section.

In another embodiment according to any of the previous embodiments, thegear reduction is positioned intermediate the fan and a compressor rotordriven by the fan drive turbine rotor.

In another embodiment according to any of the previous embodiments, thegear reduction is positioned intermediate the fan drive turbine rotorand a compressor rotor driven by the fan drive turbine rotor.

In another featured embodiment, a method of designing a gas turbineengine comprises the steps of including a first turbine rotor to drive acompressor rotor and a fan turbine rotor for driving a fan through agear reduction. A number of blades is selected in at least one row ofthe compressor rotor, in combination with a rotational speed of thecompressor rotor, such that the following formula holds true for the atleast one row of the compressor rotor: (the number of blades×therotational speed)/60 s≧5500 Hz. The rotational speed is in revolutionsper minute.

In another embodiment according to the previous embodiment, the formularesults in a number greater than or equal to about 6000 Hz.

In another embodiment according to any of the previous embodiments, thegas turbine engine is rated to produce about 15,000 pounds of thrust ormore.

In another embodiment according to any of the previous embodiments, therotational speed is an approach speed.

In another embodiment according to any of the previous embodiments, theturbine section includes a higher pressure turbine rotor and a lowerpressure turbine rotor. The fan drive turbine rotor is the lowerpressure turbine rotor.

In another embodiment according to any of the previous embodiments, thecompressor rotor is a lower pressure compressor rotor, and the higherpressure turbine rotor drives a higher pressure compressor rotor.

In another embodiment according to any of the previous embodiments, thefirst turbine rotor and the fan turbine rotor are provided by a singlerotor.

These and other features of this application will be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

FIG. 2 shows another embodiment.

FIG. 3 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown), or an intermediate spool,among other systems or features. The fan section 22 drives air along abypass flowpath B in a bypass duct defined within a nacelle 15, whilethe compressor section 24 drives air along a core flowpath C forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57further supports bearing systems 38 in the turbine section 28. The innershaft 40 and the outer shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The terms “low” and “high” as applied to speed or pressure for thespools, compressors and turbines are of course relative to each other.That is, the low speed spool operates at a lower speed than the highspeed spool, and the low pressure sections operate at lower pressurethan the high pressures sections.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient °R)/(518.7°)°R]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second.

The use of the gear reduction between the low speed spool and the fanallows an increase of speed to the low pressure compressor. In the past,the speed of the low pressure turbine and compressor has been somewhatlimited in that the fan speed cannot be unduly large. The maximum fanspeed is at its outer tip, and in larger engines, the fan diameter ismuch larger than it may be in smaller power engines. However, the use ofthe gear reduction has freed the designer from limitation on the lowpressure turbine and compressor speeds caused by a desire to not haveunduly high fan speeds.

It has been discovered that a careful design between the number ofrotating blades, and the rotational speed of the low pressure turbinecan be selected to result in noise frequencies that are less sensitiveto human hearing. The same is true for the low pressure compressor 44.

A formula has been developed as follows:

(blade count×rotational speed)/60s≧5500 Hz.

That is, the number of rotating blades in any low pressure turbinestage, multiplied by the rotational speed of the low pressure turbine 46(in revolutions per minute), divided by 60 s should be greater than orequal to 5500 Hz. The same holds true for the low pressure compressorstages. More narrowly, the amounts should be above 6000 Hz. A worker ofordinary skill in the art would recognize that the 60 s factor is tochange revolutions per minute to Hertz, or revolutions per one second.

The operational speed of the low pressure turbine 46 and low pressurecompressor 44 as utilized in the formula should correspond to the engineoperating conditions at each noise certification point defined in Part36 or the Federal Airworthiness Regulations. More particularly, therotational speed may be taken as an approach certification point asdefined in Part 36 of the Federal Airworthiness Regulations. Forpurposes of this application and its claims, the term “approach speed”equates to this certification point.

It is envisioned that all of the rows in the low pressure turbine 46meet the above formula. However, this application may also extend to lowpressure turbines wherein the majority of the blade rows in the lowpressure turbine meet the above formula, but perhaps some may not. Thesame is true for low pressure compressors, wherein all of the rows inthe low pressure compressor 44 would meet the above formula. However,the application may extend to low pressure compressors wherein only themajority of the blade rows in the low pressure compressor meet the aboveformula, but some perhaps may not.

This will result in operational noise that would be less sensitive tohuman hearing.

In embodiments, it may be that the formula can result in a range ofgreater than or equal to 5500 Hz, and moving higher. Thus, by carefullydesigning the number of blades and controlling the operational speed ofthe low pressure turbine 46 (and a worker of ordinary skill in the artwould recognize how to control this speed) one can assure that the noisefrequencies produced by the low pressure turbine are of less concern tohumans.

The same holds true for designing the number of blades and controllingthe speed of the low pressure compressor 44. Again, a worker of ordinaryskill in the art would recognize how to control the speed.

In embodiments, it may be only the low pressure turbine rotor 46, or thelow pressure compressor rotor 44 which is designed to meet the meet theabove formula. On the other hand, it is also possible to ensure thatboth the low pressure turbine 46 and low pressure compressor 44 meet theabove formula.

This invention is most applicable to jet engines rated to produce 15,000pounds of thrust or more. In this thrust range, prior art jet engineshave typically had frequency ranges of about 4000 hertz. Thus, the noiseproblems as mentioned above have existed.

Lower thrust engines (<15,000 pounds) may have operated under conditionsthat sometimes passed above the 4000 Hz number, and even approached 6000Hz, however, this has not been in combination with the gearedarchitecture, nor in the higher powered engines which have the largerfans, and thus the greater limitations on low pressure turbine or lowpressure compressor speed.

FIG. 2 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 210 is driven byan intermediate pressure turbine 212, and a second stage compressorrotor 214 is driven by a turbine rotor 216. A combustion section 218 ispositioned intermediate the compressor rotor 214 and the turbine section216.

FIG. 3 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a low pressureturbine section.

The FIGS. 2 and 3 engines may be utilized with the speed and bladefeatures disclosed above.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a fan, a turbine section having afan drive turbine, and a compressor section having a first compressor; agear reduction positioned between said fan on one side and said fandrive turbine on another side, the gear reduction including an epicyclegear train having a gear reduction ratio of greater than 2.5:1; saidfirst compressor having a number of compressor blades in at least one ofa plurality of rows of said first compressor, and said blades rotatableat least some of the time at a rotational speed in operation, and saidnumber of compressor blades in said at least one row and said rotationalspeed being such that the following formula holds true for said at leastone row of the first compressor:(said number of blades×said rotational speed)/60 sec≧5500 Hz; saidrotational speed being an approach speed in revolutions per minute,taken at an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations; wherein a pressure ratio across thefan drive turbine being greater than 5; and wherein the gas turbineengine is rated to produce 15,000 pounds of thrust or more. 2-3.(canceled)
 4. The gas turbine engine as set forth in claim 1, whereinsaid fan delivers air into a bypass duct, and a portion of air into saidcompressor section, with a bypass ratio defined as the volume of airdelivered into the bypass duct compared to the volume of air deliveredinto the compressor section, and said bypass ratio being greater than10. 5-6. (canceled)
 7. The gas turbine engine as set forth in claim 4,wherein said fan comprises at least one fan blade, with a low fanpressure ratio of less than 1.45, the low fan pressure ratio measuredacross the fan blade alone.
 8. The gas turbine engine as set forth inclaim 7, wherein said turbine section includes a higher pressure turbineand a lower pressure turbine, and said fan drive turbine being saidlower pressure turbine.
 9. The gas turbine engine as set forth in claim8, wherein said first compressor is a lower pressure compressor, andsaid higher pressure turbine drives a higher pressure compressor. 10.The gas turbine engine as set forth in claim 9, wherein the gearreduction is positioned intermediate the fan drive turbine and saidfirst compressor.
 11. The gas turbine engine as set forth in claim 7,wherein the formula does not hold true for all of the rows of said firstcompressor.
 12. The gas turbine engine as set forth in claim 77, whereinthe formula holds true for at least a majority of the rows of said firstcompressor.
 13. (canceled)
 14. The gas turbine engine as set forth inclaim 34, wherein the fan has a low corrected fan tip speed of less than1150 ft/second.
 15. The gas turbine engine as set forth in claim 14,further comprising a core flowpath and a mid-turbine frame arrangedbetween the fan drive turbine and a second turbine and having airfoilspositioned in the core flowpath, the mid-turbine frame supporting atleast one bearing system, and wherein the formula results in a numbergreater than or equal to 6000 Hz for the at least one compressor row.16. A method of designing a gas turbine engine comprising the steps of:including a turbine section having a fan turbine for driving a fanthrough a gear reduction, the gear reduction positioned between the fanon one side and the fan turbine on another side and having a gearreduction ratio of greater than 2.5:1; providing a first compressorincluding a plurality of blade rows; selecting a combination of a numberof blades in at least one row of the first compressor and a rotationalspeed of the first compressor in operation, for producing noisefrequencies that are of less concern to humans; said selecting stepincluding determining the number of blades in the at least one row ofthe first compressor and determining a rotational speed of the firstcompressor in operation such that the following formula holds true forsaid at least one row of the first compressor:(said number of blades×said rotational speed)/60 sec≧5500 Hz; saidrotational speed being an approach speed in revolutions per minute,taken at an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations; and wherein a pressure ratio acrossthe fan turbine being greater than 5, and the gas turbine engine israted to produce 15,000 pounds of thrust or more. 17-18. (canceled) 19.The method as set forth in claim 16, wherein said first compressor is alower pressure compressor, and said higher pressure turbine driving ahigher pressure compressor.
 20. The method as set forth in claim 19,wherein: the fan comprises at least one fan blade, with a low fanpressure ratio of less than 1.45, the low fan pressure ratio measuredacross the fan blade alone; and the fan delivers air into a bypass duct,and a portion of air into said first compressor, with a bypass ratiodefined as the volume of air delivered into the bypass duct compared tothe volume of air delivered into said first compressor, and said bypassratio being greater than
 10. 21. (canceled)
 22. The method as set forthin claim 20, wherein the formula does not hold true for all of the rowsof the first compressor.
 23. The method as set forth in claim 35,wherein the formula holds true for at least a majority of the rows ofthe first compressor.
 24. The method as set forth in claim 23, whereinthe formula results in a number greater than or equal to 6000 Hz for atleast one compressor row.
 25. A gas turbine engine comprising: a fan, aturbine section having a fan drive turbine, and a compressor sectionhaving a first compressor; a gear reduction positioned between said fanon one side and said fan drive turbine on another side and having a gearreduction ratio of greater than 2.5:1; said first compressor having anumber of compressor blades in at least one of a plurality of rows ofsaid first compressor, and said blades rotatable at least some of thetime at a rotational speed in operation, and said number of compressorblades in said at least one row and said rotational speed being suchthat the following formula holds true for said at least one row of thefirst compressor:5500 Hz≦(said number of blades×said rotational speed)/60 sec≦6000 Hz;said rotational speed being an approach speed in revolutions per minute,taken at an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations; and wherein a pressure ratio acrossthe fan drive turbine being greater than
 5. 26. (canceled)
 27. The gasturbine engine as set forth in claim 25, wherein the formula does nothold true for all of the rows of the first compressor.
 28. (canceled)29. The gas turbine engine as set forth in claim 25, wherein: said fandelivers air into a bypass duct, and a portion of air into saidcompressor section, with a bypass ratio defined as the volume of airdelivered into the bypass duct compared to the volume of air deliveredinto the compressor section, and said bypass ratio being greater than10.
 30. The gas turbine engine as set forth in claim 38, wherein theformula holds true for at least a majority of the rows of the firstcompressor.
 31. The gas turbine engine as set forth in claim 11, whereinthe formula holds true for at least a majority of the rows of said firstcompressor.
 32. The gas turbine engine as set forth in claim 12, whereinthe formula holds true for all of the rows of said first compressor. 33.The gas turbine engine as set forth in claim 12, wherein the formularesults in a number greater than or equal to 6000 Hz for at least one ofthe majority of compressor rows.
 34. The gas turbine engine as set forthin claim 33, wherein the formula results in a number greater than orequal to 6000 Hz for each of the majority of compressor rows.
 35. Themethod as set forth in claim 22, wherein the formula holds true for atleast a plurality of the rows of the first compressor.
 36. The method asset forth in claim 24, wherein the formula results in a number greaterthan or equal to 6000 Hz for more than one row of the first compressor.37. The gas turbine engine as set forth in claim 29, wherein the formulaholds true for more than one row of the first compressor.
 38. The gasturbine engine as set forth in claim 37, wherein the fan comprises atleast one fan blade, with a low fan pressure ratio of less than 1.45,the low fan pressure ratio measured across the fan blade alone.
 39. Thegas turbine engine as set forth in claim 30, wherein the formula holdstrue for all of the rows of said first compressor.